1.0 Field of the Invention
The present invention relates to gas turbine engines associated with aircraft propulsion and, more particularly, to a method and apparatus for damping the vibrations of external tubing segments of gas turbine engines.
2.0 Related Art
Conventional high bypass ratio turbofan engines typically include a fan, booster, high pressure compressor, combustor, high pressure turbine and low pressure turbine in serial, axial flow relationship. A portion of the air entering the engine passes through the fan, booster and high pressure compressor, being pressurized in succession by each component. This air, which compresses the primary or core air flow, then enters the combustor where the pressurized air is mixed with fuel and burned to provide a high energy gas stream. This primary or core gas stream then expands through the high pressure turbine where energy is extracted to operate the high pressure compressor which is drivingly connected to the high pressure turbine. The primary gas stream then enters the low pressure turbine where it is further expanded, with energy extracted to operate the fan and booster which are drivingly connected to the low pressure turbine. The remainder of the air which enters the engine, other than the core air flow, passes through the fan and exits the engine through a system comprising annular ducts and a discharge nozzle, thereby creating a large portion of the engine thrust.
The highest temperatures in the engine are found in the combustor and turbines. For instance, in a low pressure turbine, which includes a rotor having a plurality of circumferentially spaced rotor blades extending from a rotor disc, and a stator assembly having a stator casing which surrounds the rotor blades, the temperature of the external surface of the stator casing, may be in excess of 1000.degree. Fahrenheit during a typical takeoff condition unless cooling air is applied. The stator assembly also includes multiple stages of nozzle segments and shroud segments which are supported by internal, machined flanges of the stator casing. The nozzles include outer platforms having an interior surface which forms a portion of the outer boundary of the annular flowpath for the primary or core air flow, with the remainder of the outer boundary formed by the equivalent surface of the rotor blade outer platforms, or tip shrouds. The stator shrouds include an inner abradable surface which is disposed adjacent to and radially outward of the corresponding rotor blade tips, thereby defining a radial clearance therebetween. The rotor/stator radial clearance in the low pressure turbine, as well as that for other parts of the engine, varies during transient operation of the engine and also as a function of the various steady state operating conditions of the engine. Large rotor/stator clearances are undesirable because any air passing over the blade tips is unavailable for energy extraction and therefore the efficiency of the engine is reduced. In order to maintain stator casing integrity, by avoiding excessive temperatures, and in order to optimize rotor/stator radial operating clearances, particularly at steady state cruise conditions, large bypass ratio turbofan engines typically include a low pressure turbine cooling manifold which surrounds and is mounted to the stator casing, and which impinges cooling air at selected locations on the casing exterior surface.
The low pressure turbine cooling manifold typically receives fan discharge air from the fan bypass duct via a relatively large circumferentially extending tube on the manifold. The cooling air is then distributed throughout the manifold by a plurality of axially extending plenums and circumferentially extending and axially spaced tubing segments. The manifold is designed so that the air impinges on the external surface of the casing at locations corresponding to the internal flanges used to mount the shroud and nozzle segments, thereby permitting control of the stator flowpath surface and the radial rotor/stator operating clearances.
The temperature of the cooling air is much less than the external temperature of the casing. For instance, during a typical takeoff condition the temperature of this air is typically in the range of 100.degree.-200.degree. Fahrenheit. Due to the difference in the temperature of the casing exterior and that of the cooling air passing through the manifold, it is advantageous to construct the manifold so that a portion of the interfaces between the axially extending distributor piehums and the circumferentially extending tubes comprise slip joints, allowing thermal expansion and contraction of the tubes, while the remainder of the interfaces typically comprise brazed joints. The cooling manifold is mounted to the stator casing via brackets which transmit the mechanical vibrations of the casing to the manifold. The frequency of these vibrations can be multiples of the speeds of the low and high pressure rotor components. This vibration has been known to cause a wear problem to exist at the aforementioned manifold slip joints with visible wear occurring around the periphery of the male end fittings, on the circumferentially extending tubing segments, and around the inner surfaces of the female sleeves integral with the axially extending distributor plenums. This wear problem is independent of axial or circumferential location on the manifold. If not detected or if left uncorrected, the circumferentially extending cooling tubes become loose in the slip joint, allowing leakage of cooling air which has an adverse impact on the radial rotor/stator operating clearances and on casing structural integrity. It is conceivable that extreme conditions of wear could lead to tube breakage.
Once the wear problem has been detected, the conventional method of correcting the problem is to remove the low pressure turbine cooling manifold from the engine and return it to the appropriate engine overhaul facilities for repair or modification. This removal increases operational costs due to the manhours required to remove and replace the manifolds, which must often be accomplished while the engine is installed on the corresponding aircraft, thereby adding to the complexity of the task, and also due to the manhours and materials required to repair or replace the manifolds. Manifold removals also create the need for an inventory of rotable spare component parts to avoid loss of flight time while the manifolds are repaired, and these spare parts further increase operational costs.
Conventional means for damping and/or supporting tubes or pipes in general, as well as those for damping gas turbine engine tubes in particular, typically include one or more of the following disadvantageous characteristics: expensive manufacture; difficult and expensive assembly; inability to install with tubing installed to an engine and/or inability to install with the engine installed on an aircraft, for the applicable gas turbine engines; restrain tubing to be damped from thermal expansion and contraction due to fixed attachment to support structures or due to other restraining features. Accordingly, prior to this invention, a need existed to provide a cost effective and retrofittable method and apparatus for damping vibrations of low pressure turbine cooling manifold tubing segments, thereby improving the installed life span of low pressure turbine cooling manifolds, without requiring a redesign of the manifold or manifold mount system, and wherein the damping means could be installed without removing the manifold from the engine, thereby reducing operational costs due to the avoidance of costs associated with manifold removal and replacement with the engine installed on a corresponding aircraft.